Rotorcraft are generally provided with at least one free-turbine engine. Power is then taken from a low pressure stage of each free turbine, which stage rotates at a speed in the range 20,000 revolutions per minute (rpm) to 50,000 rpm.
Thereafter, a speed-reduction gearbox is needed to connect the free turbines to the main lift and propulsion rotor, since the speed of rotation of such a rotor lies substantially in the range 200 rpm to 400 rpm: this gearbox is the main transmission gearbox.
Thermal limits on turbine engines and torque limits on main transmission gearboxes then enable three ratings to be defined for normal utilization of a turbine engine:                a takeoff rating that can be used for five to ten minutes, corresponding to a torque level for the main transmission gearbox and heating of each turbine engine that can both be accepted for a limited length of time without significant deterioration, this is the maximum takeoff power (PMD) rating;        a maximum continuous power rating during which capabilities are not exceeded, neither the capabilities of the main transmission gearbox, nor the capabilities associated with the maximum acceptable continuous heating upstream of the high pressure blades of the first stage of each free turbine: this is the maximum continuous power (PMC) rating; and        a maximum transient power rating that is set by regulation: this is referred to as the maximum transient power (PMT) rating.        
The engine manufacturer then determines the limits of each turbine engine so as to obtain an acceptable lifetime and a guaranteed minimum power for each of the above-mentioned ratings, where the guaranteed minimum power corresponds in particular to the power that can be developed by a turbine engine that is old, i.e. a turbine engine that is at the end of its recommended maximum utilization time.
These limits are generally monitored by monitoring three parameters of the turbine engine: the speed of rotation of the engine gas generator; the driving torque; and the ejection temperature of the gas at the inlet to the free turbine of the turbine engine, these three monitored parameters being written Ng, Cm, and C45 respectively by the person skilled in the art.
Furthermore, recent turbine engines are controlled and regulated by control and regulation apparatus having an electronic regulation computer known to the person skilled in the art as a full-authority digital engine control (FADEC), that serves in particular to determine the position of the fuel supply throttle as a function firstly of a regulation loop comprising a primary loop based on maintaining the speed of the rotation of the rotorcraft rotor, and secondly a secondary loop based on a setpoint value for a piloting parameter.
A FADEC also receives signals relating firstly to the monitored parameters of the turbine engine it controls, and secondly monitored parameters concerning important members of the rotorcraft such as the speed of rotation of the main lift and propulsion rotor, for example.
The FADEC then uses a digital connection to deliver these monitored parameter values to a display system of the control and regulation apparatus, which display system is arranged in the cockpit of the aircraft.
With reference to document FR 2 749 545, the display system may include a first limitation instrument that identifies and displays a parameter that constitutes a limiting parameter, specifically that one of the monitored parameters that is the closest to its respective limit.
It should be observed that the FADEC may optionally determine which parameter is the limiting parameter, with the first limitation instrument then serving for display purposes only.
Finally, the FADEC is capable of triggering various alarms if incidents should occur, a minor failure, or a total failure of fuel regulation to the turbine engine, for example.
Furthermore, the FADEC sends information to the display system via a digital connection in the event of a monitored parameter of the turbine engine exceeding a limit that has been predetermined by the manufacturer.
Consequently, the control and regulation apparatus includes a display system, sometimes referred to as an “avionics system”, together with an electronic regulation computer, the electronic regulation computer serving simultaneously: to regulate the turbine engine; to deliver information via a digital connection to the display system for the pilot; and to monitor parameters of the turbine engine.
In addition, for reasons of safety and in order to achieve a fault occurrence rate that complies with aviation requirements, the control and regulation apparatus is also provided with sensors connected via emergency analog connections to emergency instruments, e.g. displaying the speed of rotation of the engine gas generator, the driving torque, the gas ejection temperature at the inlet to the free turbine of the engine, or indeed the speed of rotation of said free turbine.
Thus, a failure in the digital connection of the FADEC or of the display system is compensated by an emergency analog connection together with emergency instruments.
Nevertheless, that known architecture for the control and regulation apparatus makes use both of a FADEC connected via a digital connection to a display system, and of a plurality of sensors connected to emergency instruments via analog connections. It will readily be understood that such an architecture is particularly burdensome to implement, since it requires numerous analog cables to be taken all the way to the cockpit.
In addition, the cockpit needs to have enough space to receive certain emergency instruments for each turbine engine.
Finally, the magnitude and the number of analog cables presents a cost that is not negligible in terms of weight, given that the distance between the turbine engines and the rotorcraft cockpit is often considerable.